Component repair using brazed surface textured superalloy foil

ABSTRACT

A superalloy component such as a gas turbine engine blade ( 40 ) having a ceramic thermal barrier coating ( 41 ) is repaired using a textured repair foil ( 30 ). A degraded region of the thermal barrier coating is removed ( 14 ) and the underlying superalloy material surface is prepared ( 16 ) for re-coating. The repair foil is includes a layer of boron-free braze material ( 34 ) and a layer of superalloy material ( 32 ) having a textured surface ( 36 ). The foil is brazed ( 18 ) to the prepared surface during a solution heat treatment effective to homogenize the braze ( 20 ). A new area of thermal barrier coating ( 46 ) is applied over the foil with a bond that is enhanced by the texturing of the foil surface.

This application claims benefit of the 15 Mar. 2013 filing date of the U.S. provisional patent application No. 61/787,153 (attorney docket number 2013P05676US).

FIELD OF THE INVENTION

This invention relates generally to the field of materials science, and more specifically to the repair of a superalloy gas turbine engine component having a thermal barrier coating (TBC).

BACKGROUND OF THE INVENTION

The hot gas path components of gas turbine engines are often formed of superalloy materials. The term “superalloy” is used herein as it is commonly used in the art; i.e., a highly corrosion and oxidation resistant alloy that exhibits excellent mechanical strength and resistance to creep at high temperatures. Superalloys typically include a high nickel or cobalt content. Examples of superalloys include alloys sold under the trademarks and brand names Hastelloy, Inconel alloys (e.g. IN 738, IN 792, IN 939), Rene alloys (e.g. Rene N5, Rene 80, Rene 142), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-750, ECY 768, 282, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys.

Modern gas turbine engines have firing temperatures that exceed the safe operating temperatures of known superalloy materials, so components such as combustors, transitions, and early row rotating blades and stationary vanes are often further protected by a thermal barrier coating applied to the exposed surface of the superalloy material. One such thermal barrier coating system includes a metallic bond coat, such as an MCrAlY material, applied to the superalloy material and overlaid by a ceramic insulating material such as yttria stabilized zirconia (YSZ).

The thermal barrier coatings of gas turbine engine components are known to suffer degradation such as erosion, corrosion, oxidation, cracking, spallation, etc. during operation of the engine. Particular areas of particular components may be most susceptible to degradation, such as the leading edge of an airfoil, or the platform or tip of a rotating blade. Engines are periodically dismantled and inspected, and degraded components are removed for refurbishment or replacement as appropriate. A degraded area of thermal barrier coating material may be removed by mechanical or chemical means, the underlying substrate material inspected and repaired as appropriate, and a fresh thermal barrier coating applied. The component may also be subjected to a solution heat treatment in order to restore the mechanical properties of the superalloy material.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 is a flow chart illustrating steps of a method for repairing a component having a thermal barrier coating.

FIG. 2 is a partial cross-sectional view of a repair foil as may be used in the method of FIG. 1.

FIG. 3 is a perspective view of a gas turbine engine blade having a repaired leading edge, using a cut-away view to illustrate material layers of the leading edge.

FIG. 4 is a partial cross-sectional view of a repair foil including a bond coat material layer.

DETAILED DESCRIPTION OF THE INVENTION

While existing repair techniques are effective for returning some gas turbine hot gas path components to service, the repaired components remain vulnerable to the same types of degradation as had originally caused the component to degrade. The present inventors have developed a process that not only repairs a degraded component, but also can reduce the vulnerability of the repaired component to the damaging conditions experienced upon return to service. Advantageously, when the inventive process is applied to a degraded area of a component removed from service, the improvement is thus applied precisely to the most vulnerable (previously degraded) area of the component, thereby potentially extending the operational life of the repaired component to beyond that obtained by the new component. Furthermore, the inventive process can be applied during the manufacturing of a new component to extend its service life when such areas of vulnerability are known from experience on other components or from analytical predictions.

A method in accordance with one embodiment of the invention is described with reference to FIG. 1. One skilled in the art will appreciate that some of the steps illustrated in FIG. 1 are optional and may not be included in other embodiments. A component, such as a superalloy gas turbine blade, is removed from service 10 and is subjected to an inspection 12. Degraded areas of a thermal barrier coating of the component are removed 14 by any known process to expose the underlying superalloy substrate material. The exposed substrate material is prepared 16 by cleaning after any exposed and repairable flaw has been repaired. A repair foil is then attached to the prepared substrate material by brazing 18, as will be discussed more fully below. The brazing step 18 may also function to repair small discontinuities in the exposed substrate material surface as braze material flows into the discontinuities. The braze is homogenized 20 by a heat treatment that advantageously is performed simultaneously with a solution teat treatment used to restore mechanical properties of the superalloy substrate material. A new replacement thermal barrier coating is applied over the repair foil 22, and the component is returned to service 24.

An embodiment of a repair foil 30 as may be used in step 18 of FIG. 1 for one embodiment of the invention is illustrated in FIG. 2. The foil 30 may be a composite structure including a layer of alloy material 32 joined with an underlying layer of braze material 34. The foil 30 is advantageously flexible due to its thinness, such as being 0.125″ (3.175 mm) thick in one embodiment, thereby allowing it to conform to non-planar surfaces being repaired. For the repair of a superalloy gas turbine engine component, the alloy material 32 may be the same superalloy material as that of the component or a compatible superalloy material, and the braze material 34 may be any known material including a boron and silicon free braze material such as described in co-pending U.S. patent application Ser. No. 13/495,223 (attorney docket 2011P25126US01), incorporated by reference herein.

As illustrated in FIG. 2, a top surface 36 of the foil 30 opposed the braze material 34 may be textured to have surface irregularities sized and shaped to be effective to provide an improved bond with a later-applied thermal barrier coating when compared to a similar bond made to a surface without such surface irregularities. The textured surface 36 of the repair foil 30 may be formed by any known method, for example by etching, by electron beam or laser engraving, or by being cast using a process effective to form such irregularities. One such casting process is described in U.S. Pat. No. 7,411,204 B2 and related patents assigned to Mikro Systems, Inc. Alternatively, the repair foil 30 may be formed with a spark plasma sintering process wherein respective layers of powdered braze material 34 and powdered superalloy material 32 are compressed together under pressure and heat between conductive electrodes of a molding fixture while an electrical current is passed between the electrodes and through the powdered material. Localized heating occurs between adjacent particles of the powders as a result of the electrical current, and the heat and pressure are effective to sinter the particles together. The electrode in contact with the superalloy powder has its surface prepared as the mirror image of the desired textured surface 36, thereby forming the textured surface 36 on the foil 30.

FIG. 3 illustrates a gas turbine engine blade 40 having a ceramic thermal barrier coating 41 repaired by applying a repair foil 30 to a leading edge region 42 of the blade 40. FIG. 3 illustrates the leading edge 42 in a cutaway view to show the prepared superalloy substrate material 44, the overlying repair foil 30, and the finished surface of the newly applied thermal barrier coating 46. It is known that the leading edge 42 is subjected to direct impingement by the hot combustion gas and tends to degrade faster than some other areas of the blade 40. The repair foil 30 is prepared in advance and is cut to an appropriate size to cover the area of degraded thermal barrier coating that was removed from the leading edge 42. The repair foil 30 is wrapped around the leading edge 42 and may be tacked in place prior to the braze joining process. Alternatively, for airfoils having cooling holes 43 in the area to be repaired, plugs 45 may be inserted through the foil 30 and into the underlying cooling holes to secure the foil 30 to the underlying prepared substrate material 44 prior to brazing. In one embodiment, the plugs 45 may be formed of a ceramic material which prevents the braze material from entering the holes 43 during the brazing process and is subsequently removed by any known chemical or mechanical process. In another embodiment, the plugs 45 may be formed of nickel or other metal or alloy that is beneficial or at least not harmful to the superalloy substrate material 44. Such metal or alloy plugs 45 may melt during the brazing process and would then be removed by re-drilling the cooling holes 43 as necessary.

Because the foil 30 has a limited thickness, it can be brazed 18 to the substrate material 44 and then coated with the new thermal barrier coating 46 essentially as thick as the original coating material without causing any unevenness in the finished surface at the edges 48 of the underlying foil 30, thereby maintaining the aerodynamic performance of the repaired component as originally designed. As a result of the improved mechanical adhesion between the textured surface 36 and the overlying new thermal barrier coating 46, the refurbished leading edge region 42 may provide improved service performance when compared to the original blade 40 not having such a feature.

The braze material 34 may be selected to be boron and silicon free and to have a melting temperature and range below a solution heat treatment temperature used to restore the material properties of the component substrate material. When using a braze material incorporating a melting point depressant such as titanium, hafnium or zirconium or other material included in the composition of the underlying superalloy substrate material, the solution heat treatment is effective to homogenize the braze such that no discontinuity exists between the superalloy material 32 of the repair foil 30 and that of the underlying substrate 44.

A ternary alloy for such applications may have compositions within the following ranges (all compositions disclosed herein are in units of wt. %):

Cr 15-25%;

Ti 15-25%;

balance Ni.

Particular braze alloys within this group may have the following compositions: Cr 16.3%, Ti 21.2%, balance Ni; or Cr 17.2%, Ti 20.9%, balance Ni. These particular braze alloy compositions exhibit a solidus temperature of about 1,205° C. and a liquidus temperature of about 1,215° C., and thus a melting temperature range of only 10° C. As such, they may be particularly useful when brazing to Alloy 247 or Rene 80. Another braze alloy within this group has the following composition: Cr 20%, Ti 20%, Ni 60%.

Other braze alloys may have compositions within the following ranges:

Cr 12-16%;

Ti 13-16%;

Al 0-2.5%;

Co 2-4%;

W 3-5%;

Mo 0-2%;

Ta 0-2%;

balance Ni.

A particular braze alloy within this group may have the following composition: Cr 14.1%, Ti 14%, Al 2.1%, Co 3.1%, W 4.1%, Mo 1%, Ta 1%, balance Ni. This particular braze alloy composition may be particularly useful when brazing to Alloy 247.

Other braze alloys may have compositions within the following ranges:

Cr 15-18%;

Ti 10-15%;

Al 0-2.5%;

Co 2-4%;

W 3-5%;

Mo 0-2%;

Ta 0-2%;

balance Ni.

A particular braze alloy within this group may have the following composition: Cr 17.57%, Ti 13.54%, Al 2.39%, Co 3.24%, W 3.47%, Mo 1.15%, Ta 0.83%, balance Ni. This particular braze alloy composition exhibits a solidus temperature of about 1,205° C. and a liquidus temperature of about 1,220° C., and thus a melting temperature range of only 15° C. As such, it may be particularly useful when brazing to Alloy 247 or Rene 80.

Other braze alloys may have compositions within the following ranges:

Cr 15-19%;

Ti 8-10%;

Al 0-2.5%;

Co 14-18%;

Mo 12-16%;

balance Ni.

A particular braze alloy within this group may have the following composition: Cr 15.12%, Ti 10%, Al 2.12%, Co 15.8%, Mo 12.97%, balance Ni. This particular braze alloy composition exhibits a solidus temperature of about 1,205° C. and a liquidus temperature of about 1,223° C., and thus a melting temperature range of only 18° C. As such, it may be particularly useful when brazing to Alloy 247 or IN 939.

A typical solution heat treatment effective to homogenize a braze joint of such alloys may be:

-   -   heat the assembly to 1,472° F. at 15-30° F. per minute;     -   hold at 1,472° F. for 20 minutes;     -   heat to 2,125° F. at 15-30° F. per minute;     -   hold at 2,125° F. for 20 minutes;     -   heat to 2,192-2,282° F. at 1-30° F. per minute;     -   hold at 2,192-2,282° F. for 2-12 hours;     -   furnace cool to 2,120-2,192° F.;     -   hold at 2,120-2,192° F. up to 20 minutes;     -   argon cool to room temperature.

FIG. 4 illustrates another embodiment of a repair foil 50 having a layer of bond coat material 52 such as an MCrAlY material disposed over the superalloy material 54 and braze material 56. In this embodiment the textured surface 58 is formed on the top surface of the bond coat material 52, thereby providing a good basis for mechanical attachment to a later-applied ceramic insulating layer. In various embodiments the braze material 34, 56 may be no more than 200 μm thick, the superalloy material 32, 54 may be 200-300 μm thick, and the bond coat material 52 may be 125-300 μm thick.

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. 

1.-20. (canceled)
 21. A foil comprising a layer of alloy material; and a layer of braze material on a side of the alloy material; wherein the layer of alloy material comprises a layer of superalloy material adjacent the braze material and a layer of bond coat material on the layer of superalloy material.
 22. A gas turbine engine component comprising the foil of claim 21, and a layer of thermal barrier coating material disposed over the layer of bond coat material.
 23. A foil comprising a layer of alloy material; and a layer of braze material on a side of the alloy material; wherein the braze material comprises a ternary alloy comprising a composition in wt. % of: Cr 15-25%; Ti 15-25%; balance Ni.
 24. A gas turbine engine component comprising the foil of claim 23, and a layer of thermal barrier coating material disposed over the layer of alloy material.
 25. A foil comprising a layer of alloy material; and a layer of braze material on a side of the alloy material; wherein the braze material comprises a composition in wt. % within the following ranges: Cr 12-16%; Ti 13-16%; Al 0-2.5%; Co 2-4%; W 3-5%; Mo 0-2%; Ta 0-2%; balance Ni.
 26. A gas turbine engine component comprising the foil of claim 25, and a layer of thermal barrier coating material disposed over the layer of alloy material.
 27. A foil comprising a layer of alloy material; and a layer of braze material on a side of the alloy material; wherein the braze material comprises a composition in wt. % within the following ranges: Cr 15-18%; Ti 10-15%; Al 0-2.5%; Co 2-4%; W 3-5%; Mo 0-2%; Ta 0-2%; balance Ni.
 28. A gas turbine engine component comprising the foil of claim 27, and a layer of thermal barrier coating material disposed over the layer of alloy material.
 29. A foil comprising a layer of alloy material; and a layer of braze material on a side of the alloy material; wherein the braze material comprises a composition in wt. % within the following ranges: Cr 15-19%; Ti 8-10%; Al 0-2.5%; Co 14-18%; Mo 12-16%; balance Ni.
 30. A gas turbine engine component comprising the foil of claim 29, and a layer of thermal barrier coating material disposed over the layer of alloy material. 